To allow the use of the diverse mission instruments of a satellite, such as telecommunications systems, it is required to be able to control the position and the orientation of the satellite in orbit. Accordingly, control systems are implemented to maintain on the one hand the orientation of the satellite with respect to the earth, this being attitude control, and on the other hand its position in orbit with respect to a desired ideal position, this being orbital control. For example in the case of geostationary satellites, orbital control seeks to limit the inclination with respect to the equatorial plane, to limit the eccentricity of the orbit, and to limit the drift of the longitudinal position of the satellite with respect to the earth. Accordingly, propulsion units are positioned at diverse locations on the satellite so as to correct the trajectory at more or less close intervals by applying a force to the satellite. These station-keeping operations make it necessary to be able to have a sufficient reserve of fuel throughout the life of the satellite.
A satellite is placed in orbit through the combination of a launcher space vehicle and of its own propulsion systems. The launcher transports and releases the satellite on a first so-called terrestrial transfer orbit, whose perigee is generally low; once on this first orbit, a propulsion system of the satellite takes over to transport the satellite to its final orbit. Generally, this transfer is carried out by means of a principal satellite propulsion unit PSP consuming a chemical fuel of propellant type, delivering a high-power thrust making it possible to rapidly reach the final orbit.
Once placed on station, several lower-power secondary propulsion units maintain the satellite in position in the orbit. Accordingly, propellant-based chemical propulsion units or electric propulsion units may be used. In an electric propulsion unit, of plasma propulsion unit or ion propulsion unit type, xenon atoms are ionized by collision with electrons, creating xenon ions. The thrust is generated when the charged xenon ions are accelerated out of the propulsion unit by an electromagnetic field. Although expensive and of significant initial mass, the effectiveness of the propulsion unit, or its capacity to generate force by ejecting mass, also called specific impulse, is substantially more significant than that of chemical propulsion units.
In the known systems, chemical propulsion units and electric propulsion units are positioned at several locations on the structure of the satellite so as to address all of the mission requirements, from transport from the transfer orbit to the maintaining of the satellite in orbit throughout its life. The drawback of the propulsion systems thus implemented is the high cost and high mass of the various propulsion units and the fuel. These drawbacks limit the satellite's payload stowage capacity.
According to the known state of the art, an orbital control system seeks to steer the position of the satellite via six orbital parameters. FIG. 1 represents a geostationary satellite 10 in orbit 11 around the earth 12. The orbit 11 is inclined by an angle θ with respect to the equatorial plane 13 which contains the ideal geostationary orbit 14. The satellite's orbit 11 cuts the equatorial plane 13 at two points 15 and 16, customarily called orbital nodes. The six orbital parameters used to describe the position of a satellite are also known: the semi-major axis, the eccentricity, the inclination, the argument of the ascending node, the argument of the perigee, and the true anomaly. Orbital control consists in quantifying these orbital parameters and in carrying out the operations necessary by means of the onboard propulsion systems, to maintain the satellite in a predefined zone around an ideal position. By way of example, for a geostationary satellite, a drift window of plus or minus 0.1°, representing a width of almost 150 km, is allotted around a target position.
A contemporary architecture, such as represented in FIG. 2, of a satellite 10 comprises a parallelepipedal structure 20 on which are fixed diverse devices useful for the steering of the satellite 10 and for its mission. Telecommunications instruments 21 are installed on a face 22 whose orientation is maintained towards the earth, commonly called the earth face. On an opposite face 23, commonly called the anti-earth face, is positioned the principal satellite propulsion unit PSP which ensures notably the thrust necessary for transfer from the low orbit to the final orbit. On two opposite lateral faces 24 and 25, commonly called the North face and the South face, because of their orientation with respect to the equatorial plane, are positioned two sets of solar panels 26 and 27 allowing the supply of electrical power to the onboard systems. Diverse devices may be carried onboard the lateral faces 28 and 29, commonly called East and West face for their orientation with respect to a terrestrial longitude. The maintaining of a constant orientation of the satellite with respect to the earth is necessary for the proper progress of the satellite mission, for example in respect of the orientation of the solar panels 26 and 27 or the pointing of the telecommunications systems 21 towards earth. This is carried out by means of an attitude control system. Several attitude control systems able to detect and correct orientation errors are known. Thus, the measurement of the satellite orientation can be carried out by means of a sensor assembly, comprising for example an earth-ward directed sensor, positioned on the earth face for a measurement in regard to two axes, pitch and roll, with respect to the earth and an assembly 30 of gyroscopes for detection of the rotation speeds in regard to three axes. On the basis of these measurements, corrections of orientation of the satellite around its centre of gravity can be made, for example by means of an assembly of inertia wheels 31 or of gyroscopic actuators.
A satellite equipped with such a system allowing attitude control is said to be stabilized in regard to three axes. Typically, by controlling the rotation speed and the orientation of the inertia wheels, one knows how to correct an orientation error in a reference trihedron tied to the satellite. Hereinafter, we call Z an axis directed towards the earth, also called the yaw axis, Y an axis perpendicular to the orbit and oriented in the sense opposite to the angular momentum of the orbit (Southwards for a geostationary), also called the pitch axis, and X an axis forming with Y and Z a right-handed orthogonal frame, also called the roll axis which is oriented along the speed in the case of circular orbits.
For orbital control, several propulsion units are disposed on the structure 20 of the satellite 10. A first propulsion unit of large power PSP, making it possible to ensure the transfer between the initial terrestrial orbit (after launcher release) and the final orbit, is positioned on the anti-earth face 23. According to a known state of the art, a first assembly of propulsion units, comprising for example two propulsion units 32 and 33 positioned in the North face and in the South face in proximity to the anti-earth face, is used to control the inclination. A second assembly of propulsion units, such as for example the propulsion units 34 and 35, positioned in the East and West faces, is used for the control of the eccentricity and the drift. It is also known that the control of the inclination requires of the order of five to ten times as much fuel as the control of the eccentricity and drift. For this reason, inclination control is in general carried out by means of plasma propulsion unit, a more frugal consumer of fuel, while the propulsion units dedicated to the control of the eccentricity and drift are usually chemical propellant based.
By way of example, a contemporary satellite of dry mass 2500 kg and making it possible to carry an onboard payload of 900 kg, comprises a principal propulsion unit, two plasma propulsion units for the inclination and the eccentricity, and four propellant-based propulsion units for the eccentricity and the drift. Typically, 1700 kg of propellant are necessary for the initial transfer of orbit, and 220 kg of Xenon are necessary to ensure the orbital control of the satellite for a mission duration of about 15 years. Thus, the cost and the mass of current propulsion systems limit the capacity to carry a heavy payload onboard. Let us also note that in most known propulsion systems for orbital control, the various onboard propulsion units comprise in reality two propulsive motors positioned side by side, for mission safety and reliability reasons. This redundancy, well known to the person skilled in the art, is not represented in the figures but it is considered hereinafter that a propulsion unit may consist of one or more propulsive motors forming a propulsive assembly, and whose deliverable thrust is identical, in orientation or in intensity.
FIGS. 3a, 3b and 3c illustrate the principle of the orbital control for a satellite according to the known state of the art. The structure 20 of the satellite 10 is represented in side view, the East face being visible. The propulsion unit 32 is linked to the North face of the structure 20 by means of a two-axis mechanism 40. The two-axis mechanism 40 allows the rotation of the propulsion unit 32 with respect to the structure 20 according to a first axis parallel to the Y axis and a second axis parallel to the X axis. In FIGS. 3a to 3c, the two-axis mechanism 40 is a gimbal link achieved by means of a first pivot link 41 of axis parallel to the Y axis and a second pivot link 42 of axis parallel to the X axis. The centre of mass of the satellite, situated inside the parallelepipedal structure 20, is referenced CM.
In FIG. 3a, the orientation of the propulsion unit 32 makes it possible to direct the thrust of the propulsion unit towards the centre of mass CM of the satellite. To perform an inclination correction manoeuvre, a technique known to the person skilled in the art consists in firing the propulsion unit 32 a first time in proximity to an orbital node, for example 15, and then the propulsion unit on the opposite side a second time in proximity to the opposite orbital node, 16 in the example. Thus, the thrust, oriented towards the centre of mass CM, of the first firing of the propulsion unit 32 displaces the satellite in a direction having a Z component and a Y component. Twelve hours afterwards, the thrust of the second firing at the opposite orbital node, displaces the satellite in a direction having a Z component opposite to the first firing, and which compensates the undesired effect thereof on the eccentricity and a likewise opposite Y component but whose desired effects in terms of inclination are compounded. Thus, two firings of equal intensities carried out at twelve hour intervals in proximity to the orbital nodes 15 and 16 make it possible to cancel the effect of the radial component and preserve only a North-South correction. This known procedure allows daily correction of the inclination.
With this same technique it is also possible, by applying a second thrust of different intensity to the first, to apply eccentricity corrections along an axis perpendicular to the line joining the two orbital nodes 15 and 16. Techniques have also been developed to allow eccentricity corrections according to a second axis, by shifting the firing of the propulsion unit with respect to the orbital node, but at the price of less good effectiveness of the control of the inclination. To summarize, the known systems make it possible by means of two propulsion unit systems 32 and 33 to ensure the control of the inclination and the control of the eccentricity along an axis without deoptimization of the inclination control, or to ensure the control of the inclination and the control of the eccentricity according to two axes with deoptimization of the inclination control. The control of the drift may not be carried out by these two propulsion units. A contemporary satellite accordingly comprises four chemical-propellant nozzles positioned on the East and West faces of the satellite.
The propulsion unit systems 32 and 33 are also useful for managing the momentum of attitude control systems, as illustrated in FIGS. 3b and 3c. By applying a thrust off the centre of mass CM—in a plane Y-Z in FIG. 3b and off the plane Y-Z in FIG. 3c, a rotation torque is generated on the satellite—a roll torque in FIG. 3b and a pitch and yaw torque in FIG. 3c. These two torques can be used to charge or discharge the inertia wheels in relation to two axes. For example, when the rotation speed of an inertia wheel reaches its limit speed, it will be sought to intentionally orient the thrust off the centre of mass CM so as to generate, in addition to the desired displacement of the satellite, a torque making it possible to desaturate the inertia wheel, or more generally, the problem will be anticipated by bringing the angular momentum down to desired values upon each manoeuvre. These desired values being able of course to be zero, but also a judiciously defined value such as to anticipate the evolution of the angular momentum between two manoeuvres under the effect of the radiation pressure, notably solar.
Let us also note that the centre of mass of the satellite varies in the course of the life of the satellite, notably because of the progressive consumption of the onboard fuel. In the known systems, algorithms are implemented for the combined management of attitude control and of orbital control, and to make it possible to take account of the position of the centre of mass CM throughout the life of the satellite.
The issue of being able to have effective propulsion systems is therefore understood. The current solutions, which implement different nature propulsion units at diverse locations of the satellite, are relatively complex and expensive, and exhibit a high mass which limits the satellite's stowage capacity.